RLV-TDStructural Model has been realised, which consists of fuselage nose body, fuselage straight body, a pair of double delta wings and two vertical tails. This structural assembly incorporates all the complexities of an aircraft and rocket embedded in it.

Qualification model of Radar Altimeter was realised and balloon test conducted at TIFR, Hyderabad. Carbon-carbon (C/C) laminates for nose cap were realised through a new route. Functional qualification test of Launch Hold and Release System (LHRS) with dual pyro initiation carried out with simulated interfaces.

The testing of HS9 booster stage separation system along with hydraulic line separation system was completed successfully. High altitude test of the 2 kN retro rocket developed for jettisoning spent HS9 motor was successfully conducted at SDSC SHAR.

The Integrated Technical Review (ITR) of RLV-TD by the National Review Committee in October 2012 has concluded that launch of RLV-TD HEX-01 mission in September 2013 is feasible

Iron Bird facility: This facility is second of its kind in the countryto simulate the actual flight profiles using the actuators, control electronics, the entire NGC hardware with built in NGCsoftware. The actual flight hydraulic lines will be truly represented and the control surface actuator movements also willbe simulated. This facility is used to carry out the Actuator in Loop Simulation runs for RLV-TD.

I read the update about a new manufacturing route for the nose cap CC laminates, but any news on the TPS testing? Actually, any details of the TPS at all? How re-usable is it intended to be?

The flight profile of the TSTO is also intriguing.1) Why is the 1st stage the winged booster, and the upper stage the grasshopper? Why not vice versa? Some benefit of the aerodynamic lift to the curvilinear trajectory followed in ascent?2) Why is there a different flyback mechanism for the upper stage, and the lower stage booster in any case? As it is, the diagram shows the first stage reaching altitudes+velocities where thermal protection becomes necessary for re-entry (unlike the recoverable, and reusable shuttle SRBs). Why not add a smaller version of the first stage on top of it instead of a grasshopper? I would think that two different technologies leads to increased design and development (if not operational also) costs.

Among those two, I'm biased to winged flyback over a vertical landing: cross range ability, possibly cheaper (propellant wise) control during ascent, and... well, they look the part of flying machines too

Wow! That's awesome antriksh. That looks very similar to what NASA was looking at for the cancelled Space Launch Initiative.

Thanks!! Yes the first stage is similar to the RBS concept. TSTO Hopes to achieve 100 flights per vehicle, semi-cryo engine design is targeting reputability of 15 flights. turn around time of one month.

I read the update about a new manufacturing route for the nose cap CC laminates, but any news on the TPS testing? Actually, any details of the TPS at all? How re-usable is it intended to be?

The flight profile of the TSTO is also intriguing.1) Why is the 1st stage the winged booster, and the upper stage the grasshopper? Why not vice versa? Some benefit of the aerodynamic lift to the curvilinear trajectory followed in ascent?2) Why is there a different flyback mechanism for the upper stage, and the lower stage booster in any case? As it is, the diagram shows the first stage reaching altitudes+velocities where thermal protection becomes necessary for re-entry (unlike the recoverable, and reusable shuttle SRBs). Why not add a smaller version of the first stage on top of it instead of a grasshopper? I would think that two different technologies leads to increased design and development (if not operational also) costs.

Among those two, I'm biased to winged flyback over a vertical landing: cross range ability, possibly cheaper (propellant wise) control during ascent, and... well, they look the part of flying machines too

For HEX01 TPS will be a combination of ablative, silica tiles, ceramic matrix and carbon composites and probably matellic. There is no data available on the reusability of the TPS, but it seems that ISRO is more inclined towards using metallic TPS for the TSTO. ex.

1. paper, "Manufacturing of Inconel 718 Based Honeycomb Panels for Metallic Thermal Protection Systems" 2012Abstract: Metallic thermal protection system (MTPS) offers significant improvements over the ceramic based TPS for reentry applications. Space shuttle refurbishment time is estimated to be around 17000 man hours between flights. Metallic based TPS can be fabricated easily and provides wide range of design options for TPS. Adaptability and robustness of metallic thermal protection systems offers the potential for reusability. In this work, a unique manufacturing process has been evolved to realize light weight honeycomb panels through corrugation, laser welding and diffusion brazing of faceplates, where in 50 micron thick Inconel718 foil is used for making honeycomb core and 0.2mm thick Inconel718 foil as faceplates. The compression and three point bend test on these panels have shown no debond between faceplates and honeycomb core. 150x150x5mm size honeycomb panels were coated with YSZ and NiCrAlY based Thermal Barrier Coatings (TBC) and high temperature tests have shown thermal resistance of around 570 0C with front wall temperature of 1186 0C and back wall of 533 0C. Also these panels have been characterized for reusability by the testing of same panel at different heat flux levels. Though it is found that honeycomb panel has shown its integrity without debond a certain acceptable level of degradation in coating is observed. Thus Inconel718 based honeycomb panels with TBC coating are proved for use as thermal protection system for reusable launch vehicle systems.

Abstract: Functionally graded coating material (FGM) based on yttria-stabilized zirconia (YSZ) and Ni-Cr-Al-Y was designed and developed for metallic thermal protection system of reusable launch vehicle (RLV). Coating was made using premixed mechanically alloyed YSZ and Ni-Cr-Al-Y powders through plasma spray technique. Thermal stress analysis was carried out, which showed significant reduction in stress in FGM coating as compared to dual coating. The phase composition of coating was found to be close to the designed one. Porosity varied in the range of 8-18%. Average emissivity of three different time exposures of 30, 60 and 90 s was found to be 0.8. Solar absorptivity was found to be 0.55. Fatigue life of FGM coating evaluated along with Inconel and Ti6Al4V metallic substrate was compared with dual coating. FGM coating could be fatigue tested to relatively higher thermal cycles as compared to dual coating on the Inconel substrate. Heat flux measured at top surface was found to be close to simulated heat flux for windward side of RLV. Top surface temperature was similar for both type of metallic substrates and was matching with predicted temperature. However, substrate temperature was higher for Ti6Al4V as compared to Inconel alloy due to higher thermal diffusivity of Ti6Al4V.

3. Project : ‘Thermostructural Analysis of Metallic Thermal Protection System’: Sponsored by Vikram Sarabhai Space Centre, Thiruvananthapuram, Indian Space Research Organisation, 2008-09. Manned space missions returning to the Earth require thermal protection system to absorb the thermal energy due to aerodynamic heating. Of the four mechanisms of thermal protection: (i) heat sink, (ii) cooling, (iii) surface insulation and (iv) ablation, the third one is considered suitable for vehicles used in multiple missions, such as space shuttle. In order to avoid / reduce damage, requiring extensive repair before next flight, metallic thermal protection systems (MTPS) are considered suitable. ISRO is presently involved in the design and development of MTPS for its Re-usable Launch Vehicle (RLV) programme. A finite element method based software has been developed as a design tool to carry out thermal and structural analyses of MTPS.

4. Patent: Title of Invention MANUFACTURING PROCESS TO REALIZE LIGHTWEIGHT INCONEL-718 PANELS FOR METALLIC THERMAL PROTECTION SYSTEMAbstract ABSTRACT "A method of manufacturing lightweight, honeycomb metallic thermal protection panels." This invention relates to lightweight honeycomb metallic thermal panels, which are reusable, heat resistant and are useful in making aerospace vehicle parts. Structures made from such panels are capable of with standing temperature conditions at re-entry of space vehicles. These panels are made from honeycomb structures made from thin corrugated films of super alloys like NiCr alloy Titanium Aluminize and the like which are laser welded to form honey comb structures of the desired thickness. They are then sandwiched between two face plates, which are treated to withstand oxidation.

As far as choice of design is concerned, ISRO must have chosen the design based on required mission profile, respective aerodynamic challenges involved and over all cost analysis of different configuration. Its nothing to do with looks.

Thanks Antriksh! I remembered reading a media-piece sometime back on ISRO's design philosophy focusing on metal ablatives, especially Chromium based ones. The selection of papers you quoted shows their progress nicely. I wonder if they're investigating TBCs on pre-stressed faceplates. With those honeycomb panels, it should be interesting.

Btw, is are there some consolidated online resources for ISRO's vehicle development research that I'm not aware of? Or was that some your dedicated google-fu, crawling through several personal research pages of PIs on their various institute websites?

I read the update about a new manufacturing route for the nose cap CC laminates, but any news on the TPS testing? Actually, any details of the TPS at all? How re-usable is it intended to be?

The flight profile of the TSTO is also intriguing.1) Why is the 1st stage the winged booster, and the upper stage the grasshopper? Why not vice versa? Some benefit of the aerodynamic lift to the curvilinear trajectory followed in ascent?

First stage needs to travel back to the launch site (or some friendly landing base).

Wings help this nicely, can turn and glide aerodynamically.

Second stage can stay in orbit long enough so that it can pass over the launch site and land to same site without need for long-distance aerodynamic gliding.

Shuttle orbiter did not need those big wings for missions it was used, it had those because DoD wanted to be able to do single-orbit reconnaissance missions over soviet union from Vandenberg.

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2) Why is there a different flyback mechanism for the upper stage, and the lower stage booster in any case? As it is, the diagram shows the first stage reaching altitudes+velocities where thermal protection becomes necessary for re-entry (unlike the recoverable, and reusable shuttle SRBs). Why not add a smaller version of the first stage on top of it instead of a grasshopper? I would think that two different technologies leads to increased design and development (if not operational also) costs.

Because the those wings on second stage would add extra weight that would reduce the payload weight. And bigger thermal shield would be needed.

Also, those wings would add extra aerodynamic drag in the beginning of the ascent.

First stage may needs _some_ thermal shielding, but lighter because it's re-entering at much lower speed that the second stage.

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Among those two, I'm biased to winged flyback over a vertical landing: cross range ability, possibly cheaper (propellant wise) control during ascent, and... well, they look the part of flying machines too

If cross-range is not needed, then it's not needed.

And I don't understand your point of aerodynamic control during ascent. The first stage can do it without the second stage, so having or not having aero surfaces on the second stage has nothing to do with it.

You contend that the wings on the second stage would add drag and extra weight. But the second stage flyback booster would have to carry propellant with it too - the weight of the wings substitute that. Sure, there might be more aerodynamic drag, but you can exploit that for propellant-less steering of the vehicle on its ascent. You can maintain control with first AND second stages. It adds redundancy if nothing else. Also, it seems a lot more complicated to have to store propellant over several orbits, through the thermal and dynamic conditions of re-entry and then ensuring the functionality of an engine soon after (in contrast to relatively passive aerodynamic surface control).

As regards the different levels of thermal shielding required: that may be true, but I'm sure that there are conventional cost savings if both stages' TPS systems derive from (if not use) the same technology and architectures. The simplicity of the supply chain - from design, manufacturing, tooling, integration, to the costs of training personnel etc. ought to be grow with similarity.

I'm sure that these and other concerns were accounted for, before they selected either type. I'd just like to go through the numbers that led to this design choice is all. The only clear winner would be if we go ahead and develop a TPS for the upper stage, which can withstand a capsule-like re-entry (aided only by parachutes for the final descent) AND IS ALSO re-usable (or design a capsule that can easily take on a new set of tiles/coating). Like a reusable Soyuz crew capsule.

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Btw, I don't understand what you mean about the shuttle's wings. The orbital (as opposed to sub-orbital) shuttle will come around over friendly territory of its own accord anyway - it doesn't have to have a glide ability. You might as well do it from Kennedy. If surprise, and ASAT missiles were key concerns, you could carry additional propellant for last minute orbit change/conjunction avoidance manoeuvres - and still stay below the required mass of a retrograde Vandenburg orbital launch.

Well, I think that a nice incidental benefit of all these different programs and efforts worldwide is the potential for a diversity of design configurations and combinations, which can then offer opportunity for comparisons through real-world performance and results.

While I certainly hope that SpaceX's F9R first stage is able to execute a glorious tail-first landing -- I really hope they telecast it live -- some people on the forum have expressed fears that it may tumble out of control. Wings at least seem to be a more familiar approach, and I'm thinking a more reliable one.

A large number of RLV concepts can be derived considering features like1) number of stages, 2) partial or full re-usability, 3) vertical or horizontal take-off / landing,4) tandem or parallel staging, 5) wing body or lifting body etc.

During the last decade extensive studies have been carried out in ISRO and based on detailed trade-off assessment, it was found that two stage to orbit configuration with a semi cryogenic winged booster and a cryogenic ballistic orbiter is the most feasible option considering the near term technologies.